A typical gas turbine engine generally possesses a forward end and an aft end with its several core or propulsion components positioned axially therebetween. An air inlet or intake is located at a forward end of the gas turbine engine. Moving toward the aft end, in order, the intake is followed by a compressor, a combustion chamber, and a turbine. It will be readily apparent from those skilled in the art that additional components may also be included in the gas turbine engine, such as, for example, low-pressure and high-pressure compressors, and low-pressure and high-pressure turbines. This, however, is not an exhaustive list. The gas turbine engine also typically has an internal shaft axially disposed along a center longitudinal axis of the gas turbine engine. The internal shaft is connected to both the high pressure turbine and the high pressure compressor, such that the high pressure turbine provides a rotational input to the high pressure compressor to drive the compressor blades.
Higher operating temperatures for gas turbine engines are continuously being sought in order to improve their efficiency. However, as operating temperatures increase, the high temperature durability of the components of the gas turbine engine must correspondingly increase. Significant advances in high temperature capabilities have been achieved through the formulation of iron, nickel, and cobalt-based superalloys. While superalloys have found wide use for components used throughout gas turbine engines, and especially in the higher temperature areas, alternative lighter-weight component materials have been proposed.
Turbine shrouds and blades may be made of a number of materials, including nickel-based superalloys and ceramic matrix composites (CMCs). CMCs are a class of materials that consist of a reinforcing material surrounded by a ceramic matrix phase. Such materials, along with certain monolithic ceramics (i.e. ceramic materials without a reinforcing material), are currently being used for higher temperature applications. These ceramic materials are lightweight compared to superalloys yet can still provide strength and durability to the component made therefrom. Therefore, such materials are currently being considered for many gas turbine components used in higher temperature sections of gas turbine engines, such as airfoils (e.g. turbines, and vanes), combustors, shrouds and other like components that would benefit from the lighter-weight and higher temperature capability these materials can offer. CMCs are an attractive alternative to nickel-based superalloys for turbine applications because of their high temperature capability and light weight.
Within the high pressure and low pressure turbines, a shroud is a ring of material surrounding the rotating blades. The shroud assembly circumscribes the turbine rotor and defines an outer boundary for combustion gases flowing through the turbines. The turbine shroud may be a single unitary structure or may be formed of a plurality of segments.
Turbine performance and efficiency may be enhanced by reducing the space between the tip of the rotating blade and the stationary shroud to limit the flow of air over or around the top of the blade that would otherwise bypass the blade. This bypass causes loss of efficiency in the gas turbine engine. During engine operation, the blade tips can rub against the shroud, thereby increasing the gap and resulting in a loss of efficiency, or in some cases, damaging or destroying the blades.
For CMC shrouds, damage to metal blade is even more likely since the silicon carbide material is significantly harder than the nickel-based superalloys. For CMC shrouds, an environmental barrier coating is also required for successful performance/survival of the part due to material loss from high temperature steam recession.
In order to reduce the risk associated with coating loss, an abradable layer is deposited on top of the environmental barrier coating to protect from blade rub. It may be desirable that the abradable layer is formed of a series of ceramic ridges that break away upon contact with the rotating blade tip. The ridges are designed to break in order to inhibit damage to the blades during operation.
Abradable coatings have been applied to CMC shroud components to insure breakaway of the abradable coating instead of damaging metal blades. The abradable coatings have been applied by a plasma spray process where only a small fraction of the sprayed material is comprised in the abradable coating. Moreover, if the abradable coating is patterned using a series of abradable ridges, utilization of the material is further reduced, since the coating is sprayed onto a metal mask to only allow material through the mask to form the ridges.
As may be seen by the foregoing, it would be desirable to improve these aspects of gas turbine engine components. For example, it would be desirable to deposit an abradable coating on either of the blade or shroud which inhibits the damage to blades. It would further be desirable to deposit an abradable coating using a method that allows for significantly greater material utilization (i.e. less waste of the material being deposited) particularly since the material involved typically are comprised of at least one rare earth element.
The information included in this Background section of the specification, including any references cited herein and any description or discussion thereof, is included for technical reference purposes only and is not to be regarded subject matter by which the scope of the embodiments of the present invention is to be bound.